A gas turbine engine typically includes a gas generator, which is a core engine having in serial axial flow relationship, a high pressure compressor to compress the airflow entering the core engine, a combustor in which a mixture of fuel and the compressed air is burned to generate a propulsive gas flow, and a high pressure turbine which is rotated by the propulsive gas flow and which is connected by a large diameter shaft to drive the high pressure compressor. Rocket engines have also been proposed as gas generators for high speed applications such as that shown in U.S. Pat. No. 5,074,118. However, the use of a rocket engine as the gas generator requires that the craft carry its own oxygen supply which is a highly undesirable weight penalty. The present invention is directed at gas turbine engines incorporating gas generators which are Brayton Cycle engines having combustors that use air ingested from the atmosphere. A typical bypass turbofan engine also has a low pressure turbine aft of the gas generator and a forward fan forward of the gas generator. Where the gas generator is a core engine, a typical bypass turbofan engine adds a low pressure turbine aft of the high pressure turbine and adds a forward fan forward of the high pressure compressor.
In a typical variable bypass ratio design, as disclosed in U.S. Pat. No. 4,068,471, the front fan includes one or more forward rows of fan rotor blades connected to a small diameter drive shaft, which runs through a hollow large diameter shaft of a core gas generator and is driven by the low pressure turbine. The core gas generator includes an aft fan having one or more aft rows of fan rotor blades connected to the larger diameter drive shaft which is driven by the high pressure turbine and are disposed in serial, axial flow relationship between the forward fan and the high pressure compressor. A variable area bypass injector is located between the forward and aft fans to vary the amount of air entering a first inlet of a fan bypass duct which varies the fan bypass ratio of the engine (i.e., the ratio of the air flowing through the fan bypass duct to the air flowing through the core engine) from which comes the term variable cycle to describe the engine. The fan bypass duct has a second inlet located aft of the aft row of fan blades. Control of airflow directed into the first and second bypass duct inlets may be accomplished by selector valve mechanisms and some more particular valves called variable bypass injectors commonly referred to as VABIs.
A variable cycle aircraft gas turbine engine having a fan bypass duct with two or more inlets may be called a multiple bypass stream variable cycle gas turbine engine. A row of stator vanes is typically located just forward of each forward and aft row of fan blades. Selected rows of stator vanes are variable, typically variable angle, to vary the angle of the flow seen by the rotor blades. Some of the engine thrust comes from the propulsive gases exiting the core engine and some from the airflow exiting the fan bypass duct.
Two variable cycle aircraft gas turbine engines having a core driven supercharger in a bypass duct with a variable stator tip controlling the airflow and pressure ratio of its core driven fan tip in the bypass duct are disclosed in U.S. patent application Ser. No. 08/624,288, filed Mar. 29, 1996, entitled "TURBOFAN ENGINE WITH A CORE DRIVEN SUPERCHARGED BYPASS DUCT AND FIXED GEOMETRY NOZZLE" and in U.S. patent application Ser. No. 08/625,498, filed Mar. 29, 1996, entitled "TURBOFAN ENGINE WITH A CORE DRIVEN SUPERCHARGED BYPASS DUCT". These two references disclose multiple bypass stream variable cycle gas turbine engines capable of efficiently holding core airflow essentially constant. Alternatively, these engines are capable of varying airflow while the bypass flow and energy that are added to the second inlet of the bypass duct can be independently decreased without wasting energy in the form of higher than needed pressure in the bypass flow to the bypass duct supplied through the second inlet.
New variable cycle hyperjet (VCHJ) engines incorporating Brayton Cycle gas generators are presently being developed to power several types of aircraft from sea level takeoff conditions to Mach 5.5. One such engine utilizes turbomachinery and an augmentor to provide thrust up to approximately Mach 4.0, where the turbomachinery becomes ineffective. At Mach 4.0 the augmentor is transitioned to a hyperburner as the engine airflow is diverted around the turbomachinery which is then shut down and cocooned for thermal protection. The engine continues to operate at high Mach No. supersonic conditions as a ramjet. The gas generator may be a core engine having high pressure compressor and turbine sections and a fuel rich burning combustor therebetween. Some proposed engines do not incorporate a core engine but use some other type of Brayton Cycle engines with a very fuel rich combustor.
The fuel rich combustor limits the gas path temperature through the turbomachinery of both the core engine, if used, and the low pressure turbine section until the fan bypass air mixes with the core stream behind the low pressure turbine. The remaining fuel in the hot gas flow exiting the gas generator is ignited and combusted after at least the beginning of mixing in an exhaust duct combustor that resembles a conventional augmentor. This augmentor may or may not require any additional fuel be added in the exhaust duct combustor, thereby, possibly eliminating the need for augmentor type fuel injectors and/or conventional bluff body or other types of flame holders. The turbomachinery of the engine may be allowed to wind mill up to the max flight speed such as Mach 5.5. Fuel continues to be added through the main combustor, thus, quenching the temperature of the low pressure turbine as well as turbomachinery of the core engine if so incorporated but producing the required thrust in the secondary combustor or augmentor downstream of the mixing.
A significant performance penalty may exist in such an engine at lower speed sea level conditions where the cycle is most dependant upon excess horsepower generated by the fuel rich Brayton Cycle gas generator. A similar penalty is expected at all of the low Mach No. operating flight conditions. Engines incorporating conventional turbomachinery for a core gas generator can generally be designed to provide the required excess horse power (the net energy resulting from the gas generator which is available to drive a low pressure turbine, or provide thrust, or a combination of the two). Brayton cycle gas generators are relatively limited in their pressure rise characteristics, especially at low flight mach number conditions.
While the cycle looks acceptable at high flight Mach numbers, there is a significant mismatch in the fan and core stream exit pressures at lower speed sea level conditions which would significantly degrade the cycle of such an engine. This is because both the core and the fan bypass streams begin at the same fan discharge pressure. When the LP turbine pressure ratio required to drive the fan exceeds the net pressure rise of the Brayton cycle gas generator, the fan duct discharge pressure will exceed the LP turbine discharge pressure. However, these pressures must be equal in the engine mixer plane where the two streams come together. The excess energy in the fan duct discharge stream flow due to the pressure difference is wasted by the cycle to the extent that the bypass stream velocity cannot be used to transfer the energy to the combined stream. At high flight mach numbers, this difference is small, and has little effect on the cycle. Therefore, a great need exists to better match the fan and core stream exit pressures at lower speed sea level conditions and low Mach No. operating flight conditions of such engines. The present invention provides a variable supercharged bypass duct around the gas generator that according to one analysis can for example, at sea level take off conditions, provide an increase of over 18% in specific thrust and a decrease of about 12% in specific fuel consumption over a comparable engine without the present invention.
Though not incorporating a Brayton Cycle gas generator, U.S. Pat. No. 5,074,118 discloses an air turbo-rocket engine that has a rocket gas generator to drive a turbine which powers a supersonic fan. The fuel rich hydrogen/oxygen gas-generator supplies the correct amount of flow to drive the turbine and provide the power necessary to drive the fan. The hydrogen rich turbine exhaust flow is mixed with the main airstream in a mixing section and burned in a subsonic combustor section before being exhausted through a nozzle. Additional hydrogen may be added prior to combustion to make the overall flow stoichiometric.